Jet turbine buckets and method of making the same



B. E. KRAMER May 1, 1962 JET TURBINE BUCKETS AND METHOD OF MAKING THE SAME Filed Oct. 9, 1958 INVENTOR. 5W6! KIA/V5? M If RIVEKS United grates Patent Qfifice 3,032,315 l atented May 1, 1962 the Air Force Filed Oct. 9, 1958, Ser. No. 766,370 5 Claims. (Cl. 253-77) The present invention relates to turbine buckets and, as illustrated herein, relates more particularly to turbine buckets adapted for us at high temperatures.

Jet engines and the like are usually provided with an axial flow turbine which is operated by exhaust gases to drive a blower for furnishing air to the burners. These turbines operate at excessively high temperatures somewhat about 2000 F. Turbine buckets must have sufficient strength, toughness, creep resistance and resistance to oxidizing gases to enable the bucket to operate efiiciently without deformation or corrosion.

Molybdenum is one of the metals which exhibits high strength, toughness and creep resistance at temperatures .above 1400 F. Molybdenum, however, cannot be used at these high temperatures since the trioxide of molybdenum, which is formed under the oxidizing conditions present in a jet turbine, sublimes very rapidly at temperatures in excess of 1463 F. and a molybdenum turbine bucket will disappear in a matter of minutes.

The application of coatings to molybdenum alloys is restricted to relatively low temperatures, about 2200- F. because the molybdenum alloys recrystallize above that temperature with the loss of many desirable physical properties. This low temperature coating application precludes the use of a wide variety of refractory metals, intermetallics, cermets and ceramics; because to be effective, they require firing or sintering temperatures ranging from 2700 F. to 400G F.

The present invention contemplates the application of high temperature coatings in a manner to avoid the problems noted above. To this end, it is proposed to provide a sheath formed of molybdenum wire screen or sheet arranged to enclose molybdenum alloy turbine bucket airfoil. The molybdenum wire screen or sheet metal sheath is coated on the outside with an appropriate cermet or ceramic and then sintered at the appropriate high temperature. As disclosed herein, the molybdenum alloy turbine bucket is flame sprayed with a suitable brazing alloy which is fused on the turbine bucket airfoil. The coated and sintered molybdenum boot is then placed over the molybdenum bucket and brazed in an autoclave by usual means.

Another object of the invention is to provide a simple and effective method for providing turbine bucket airfoils with a coated boot which will Withstand relatively high temperatures. To this end, a suitable brazing alloy is flame sprayed and fused on the turbine bucket airfoil. A boot of sheet molybdenum or wire mesh is constructed to envelope completely the turbine bucket airfoil. The molybdenum boot is removed and coated on the outside with an appropriate cermet or ceramic. The coated molybdenum boot is sintered at the appropriate temperature. Finally the coated and sintered boot is placed over the molybdenum bucket and brazed in an autoclave by usual means.

With the above and other objects and features in view, the invention will now be described in connection with the accompanying drawings in which:

FIG. 1 is a view of a turbine bucket constructed according to the present invention;

FIG. 2 is a view in section taken along the lines II-II .of FIG. 1; and

FIG. 3 is an enlarged sectional view taken along the line III-III of FIG. 2.

As illustrated in FIG. 1, the bucket 10 comprises a body 12 formed of molybdenum metal or a high temperature molybdenum alloy having its bucket or airfoil surface covered with a boot 14 formed, as shown, of molybdenum wire screen. It is evident, however, that the boot 14 could be formed of a thin sheet of molybdenum metal or molybdenum alloy. It is preferred, however, to use molybdenum wire screen since it may be more readily shaped to fit the airfoil section of the molybdenum bucket body 12. The body 12 of the bucket 19 is provided, preferably, with a fir tree root 18 for securing the same in the turbine wheel.

The airfoil section of the bucket is coated by a flame spraying method by the use of a suitable brazing alloy 20. A brazing alloy containing any suitable metal such as silver solder, copper or nickel is sprayed and fused on the airfoil surface of the bucket to provide an effective means by which the boot 14 is secured to the bucket body 12. The flame spraying method is preferred but other suitable methods may be used if so desired.

After the brazing alloy coating 29 has been applied to the airfoil surface of the bucket body 12, the boot 14 is shaped to fit closely about the airfoil surface of the bucket body 12. After the boot 14 has been shaped, it is removed from the bucket body 12 and coated on the outside with a suitable ceramic having physical properties sulficient to withstand the high temperature conditions with a jet engine. A preferred corrosion resistant, shock resistant, and abrasion resistant ceramic 16 may be formed of NlCI'3Cg, (Ir-A1 0 Cr-CrB, specially compounded ceramic glass and pure oxides such as A1 0 and ZrO The selected material is preferably reduced to a fine powder and uniformly dispersed to form a slip by wet mixing.

The slip is applied to the outer surface of the boot 14 and is fired or sintered at high temperatures ranging from 2700 F. to about 4000 F. Firing or sintering at such high temperatures permits the use of a wide variety of refractory metals, intermetallics, cerrnets and ceramics which were precluded from use when the corrosion resisting coating was applied directly to the airfoil section of the turbine bucket blade.

After the boot 14 has been coated and sintered, it is placed over the molybdenum turbine bucket body 12. The assembly is then brazed in an autoclave by usual means to fuse the coated boot 14 to the molybdenum tur= bine bucket body 12.

The above method of forming a coated turbine bucket blade presents many important advantages over prior methods. The use of a molybdenum boot formed of wire screen presents additional advantages since better bonding between the bucket and the boot 14 results from the flow of brazing alloy around individual wires of the mesh or screen. Further, the use of a screen provides a boot which compensates for thermal expansion differences by providing a coating with sufficient ability to absorb thermal stress. In addition, the screen may be readily shaped about the turbine bucket and provides a ductile metallic matrix in which the refractory metals and ceramics can be impregnated.

The present invention presents the advantages to a greater degree when Wire screen is used but they are also present, although to a lesser degree, when molybdenum sheet metal is used to form the boot. In any event, the fact that the coating is produced or constructed apart from the molybdenum turbine bucket permits sintering of the coating at extremely high temperatures without harm to the turbine bucket. As a result, molybdenum or molybdenum alloy turbine buckets can be coated with materials having superior high temperature strength, oxidation resistance and abrasion resistance.

Having thus described my invention what I claim as new and desire to secure by Letters Patent of the United States is:

1. A unitary composite vane for a jet turbine comprising a root section and an airfoil blade section formed of a high temperature high strength metal, a hollow cover member open at one end for encompassing only said blade section, said cover comprising a sheath of high temperature high strength metal, and an oxidation resistant coating which requires sintering at a temperature which is harmful to said metals for said sheath, said coated sheath being bonded to said blade section for protecting said blade section against oxidation and high temperaturecorrosion.

2. A unitary'composite vane for a jet engine comprising a root section and an airfoil blade section having a core formed of a high temperature high strength molybdenum alloy, a hollow, high temperature, high strength sheath open at one end for encompassing only said blade section, said sheath comprising a high temperature high strength alloy having an outer coating of oxidation resistant material which requires sintering at a temperature which would cause a recrystallization and a reduction in the desirable physical properties of said alloys, and means for bonding said coated sheath to said blade section core.

3. A unitary composite vane for a jet engine compris ing a root section and an airfoil section having a core of high temperature high strength molybdenum alloy, a hollow woven sheath open at one end for encompassing only said blade section, said sheath comprising a high temperature high strength molybdenum sheet having an outer ceramic coating which requires sintering at a temperature which would cause recrystallization of said molybdenum alloy for protecting said molybdenum core against oxidation and high temperature corrosion, and means for bonding said coated molybdenum sheath to said core.

4. A unitary composite vane for a jet engine comprising a root section and an airfoil section having a core of high temperature high strength molybdenum alloy, a hollow sheath open at one end for encompassing said blade section, said sheath comprising a high temperature high strength molybdenum wire screen having an outer ceramic coating which requires sintering at a temperature which would cause recrystallization of said molybdenumv alloy for protecting said molybdenum core against oxida tion and high temperature corrosion, and means for bond-- References Cited in the file of this patent UNITED STATES PATENTS 2,431,660 Gaudenzi Nov. 25, 1947 2,711,973 Wainer et al. June 28, 1955 2,763,919 Kempe et al Sept. 25, 1955 2,783,966 Sorensen Mar. 5, 1957 2,819,515 Roush Jan. 14, 1958 2,856,675 Hansen Oct. 21, 1958 FOREIGN PATENTS 731,161 Great Britain June 1, 1955 1,073,330 France Mar. 24, 1954 OTHER REFERENCES Publication: NACA RM E51H23, Wire Cloth as Porous Material For Transpiration-Cooled Walls by E. R. G. Eckert, Martin R. Kinsler and Reeves P. Cochran; November 13, 1951. 

